The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel for generating hot combustion gases in a combustor. The hot gases are discharged from the combustor into a high pressure turbine which extracts energy therefrom for powering the compressor.
A low pressure turbine follows the high pressure turbine for extracting additional energy from the combustion gases for producing useful work. In a typical turbofan aircraft engine application, the low pressure turbine powers a fan disposed upstream from the compressor for producing propulsion thrust to power an aircraft. In marine and industrial applications, the low pressure turbine is joined to an output drive shaft for powering an electrical generator or propulsion screws in a ship.
The high pressure turbine may have one or more stages of stationary nozzle vanes and rotary blades, with the low pressure turbine typically including several stages of nozzles and blades. The turbine blades typically increase in size in the downstream direction as the combustion gases expand, and the temperature of the gases decreases as energy is extracted.
In view of the high temperature of the combustion gases, engine components subjected thereto typically require cooling for enhancing the life thereof. Accordingly, a portion of the air pressurized in the compressor may be channeled to various engine components for providing cooling thereof in various manners. The prior art is replete with various configurations for cooling combustor liners, nozzle vanes, rotor blades, and their associated components.
However, the air diverted for cooling the engine components is not used in the combustion process and therefore decreases engine efficiency. The known cooling configurations therefore attempt to maximize the cooling effectiveness of the diverted compressor air, which is typically used multiple times prior to being reintroduced into the exhaust path. Correspondingly, state-of-the-art superalloy materials are typically used in the turbine components for their enhanced strength at high temperature and long life. Oxidation resistance thereof is further enhanced by suitable coatings such as platinum-aluminide which further increase the durability and life of the components.
Since the combustion gases are hottest inside the combustor, the first stage high pressure turbine nozzle disposed at the outlet of the combustor requires maximum cooling effectiveness for long life. The first stage nozzle typically uses the highest pressure compressor discharge air for cooling thereof, with elaborate cooling configurations of the nozzle vanes themselves. The vanes typically have multiple internal passages for circulating the air coolant, and internal impingement baffles are typically used for impingement cooling the internal surfaces of the vanes.
The vanes typically include several rows of film cooling holes extending through the pressure and suction sides thereof which discharge the spent impingement air into corresponding films of cooling air over the external surfaces of the vane airfoil.
The pressure side of the vane airfoil is generally concave and the opposite suction side of the airfoil is generally convex, with a generally crescent shape between the leading and trailing edges of the airfoil for efficiently directing the combustion gases to the first stage high pressure turbine rotor blades. Both the temperature distribution and pressure distribution of the combustion gases over the nozzle vanes varies from the leading to trailing edges thereof, and the cooling configuration must be specifically adapted for providing balanced cooling of the nozzle vane while maintaining acceptable backflow margin. The internal pressure of the coolant in the vanes must be locally higher than the external pressure of the combustion gases to prevent backflow of the combustion gases into the film cooling holes.
The first stage rotor blades extend radially outwardly from the perimeter of a rotor disk and require correspondingly sophisticated cooling configurations different than those used for the stationary turbine nozzle. Compressor discharge air is typically used for cooling the first stage turbine blades, without discrete impingement baffles therein in view of the substantial centrifugal forces generated in the rotating blade during operation.
In a two stage high pressure turbine, a second stage turbine nozzle and second stage rotor blades are employed and typically require corresponding cooling thereof in configurations different than those for the first stage nozzle and blades in view of the different pressure and temperature distribution thereover.
The multistage low pressure turbine includes additional rows of nozzles and rotor blades which may require cooling or not depending upon the particular engine configuration. Since the combustion gas temperature is substantially reduced in the low pressure turbine, the additional complexity and need for internal cooling of the nozzle vanes and blades is typically not required.
A particular problem in cooling the low pressure turbine nozzle is the decreasing pressure distribution of the combustion gases flowing therethrough. Whereas compressor discharge air may be used for cooling the first stage turbine nozzle while maintaining acceptable backflow margins at the various rows of film cooling holes between the leading and trailing edges of the vanes, the high pressure compressor discharge air can provide excessive backflow margins when used in the low pressure turbine nozzle in view of the substantial reduction in pressure of the combustion gases.
Accordingly, one embodiment of a low pressure turbine nozzle used publicly for many years in this country bifurcates the cooling channels of the nozzle vane in two portions corresponding with the leading edge and trailing edge regions of the vane. The leading edge cooling circuit is joined in flow communication with an eight intermediate stage of the compressor, whereas the trailing edge circuit of the vane is joined in flow communication with cooling air recouped from the high pressure turbine. The recoup air has a different temperature and different pressure than the intermediate stage compressor air, and the vanes are imperforate without any outlet holes in the pressure and suction sides thereof.
In this conventional embodiment, the low pressure turbine nozzle vanes may be otherwise imperforate, with the two sources of cooling air being discharged through the inner band thereof for providing purge cooling of various forward and aft cavities found therebelow.
Marine and industrial gas turbine engines are typically derived from aircraft turbofan engines in view of the substantial sophistication and development cost thereof. The core engine including the compressor, combustor, and high pressure turbine of the turbofan engine may be used with little or no changes in the derivative marine or industrial engine. The low pressure turbine may be suitably modified with an output drive shaft for powering an electrical generator or the propulsion mechanism for a ship. However, the cooling configuration for the turbine nozzles and blades may remain unchanged in the derivative engine.
In the continuing development of derivative engines, the fan of the parent turbofan engine may be replaced by a multistage low pressure compressor driven by a new intermediate power turbine located between the high pressure turbine and the low pressure turbine. The intermediate power turbine in one configuration may use two stages of nozzles and blades.
Since the intermediate stages are located between the high pressure turbine and the low pressure turbine they are subject to the transition in pressure and temperature distribution therebetween. Since the first stage of the intermediate power turbine is disposed immediately downstream of the high pressure turbine it requires suitable cooling for the intended life.
However, the second stage nozzle of the intermediate power turbine is located downstream therefrom and immediately upstream of the low pressure turbine and does not require internal cooling of the vanes, which may therefore be simply made solid.
The first stage intermediate nozzle may be formed of a suitable superalloy, such as the same nickel-based superalloy used for the high pressure turbine nozzles, with a corresponding oxidation resistant coating such as platinum-aluminide. These high strength nozzle vanes have an associated maximum allowable metal temperature which is slightly below the temperature of the combustion gases in the intermediate power turbine.
Accordingly, the first stage nozzle of the intermediate power turbine requires additional cooling for achieving the desired life thereof, but that cooling must be effected in a new configuration being simpler and less expensive than those employed for the high pressure turbine. And, minimal additional air should be diverted from the compressor for nozzle cooling, while maintaining acceptable backflow margins.
It is therefore desired to provide a new turbine nozzle specifically configured for the operating environment of an intermediate power turbine between high and low pressure turbines.